Systems and methods for a gas turbine combustor having a bleed duct

ABSTRACT

A gas turbine system includes a compressor operative to output an airstream and a diffuser having an inlet to receive the airstream and an outlet to output the airstream. The outlet has an area larger than the inlet to diffuse the airstream. The gas turbine system also includes a fuel nozzle operative to receive fuel and emit the fuel in a combustor and at least one bleed duct having an inlet between the compressor and the outlet of the diffuser. The at least one bleed duct is operative to direct bleed air from downstream of the compressor to the fuel nozzle.

BACKGROUND OF THE INVENTION

The subject matter disclosed herein relates to gas turbine combustorsand diffusers.

Gas turbines typically include a diffuser that decelerates the airemitted from the compressor prior to the air entering the combustor toreduce combustion system pressure loss and improve engine efficiency.Packaging considerations including engine size, weight, and cost oftenresult in the optimum diffusers having relatively short lengths. Somediffusers achieve a short length by bleeding air from the air streamnear the diffuser throat to energize the air flow near the diffuser walland prevent separation of the flow from the wall and aerodynamicinstability.

BRIEF DESCRIPTION OF THE INVENTION

According to one aspect of the invention, a gas turbine systemcomprising, a diffuser operative to diffuse an airstream output from acompressor, a fuel nozzle operative to receive fuel and emit the fuel ina combustor, and at least one bleed duct operative to direct bleed airfrom down stream of the combustor to the fuel nozzle.

According to another aspect of the invention, a method for routing bleedair comprises outputting an airstream of compressed air from acompressor, drawing bleed air from the airstream down stream from thecombustor, and directing the bleed air through a duct to a fuel nozzle.

These and other advantages and features will become more apparent fromthe following description taken in conjunction with the drawings.

BRIEF DESCRIPTION OF THE DRAWING

The subject matter, which is regarded as the invention, is particularlypointed out and distinctly claimed in the claims at the conclusion ofthe specification.

The foregoing and other features, and advantages of the invention areapparent from the following detailed description taken in conjunctionwith the accompanying drawings in which:

FIG. 1 is a side partially cut-away view of a portion of a gas turbineengine.

FIG. 2 is a front partially-cut away view of the gas turbine enginealong the line A-A of FIG. 1.

The detailed description explains embodiments of the invention, togetherwith advantages and features, by way of example with reference to thedrawings.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 illustrates a side partially cut-away view of a portion of a gasturbine engine 100. The gas turbine engine 100 includes a compressorportion 102, an outer casing 104, a diffuser portion 106, a transitionpiece 108, an impingement sleeve 110, a mount portion 112 connected tothe outer casing 104, a bracket portion 114 connected to the mountportion 112, a head end volume 116 partially defined by the transitionpiece 108, a plurality of fuel nozzles 118 communicative with the headend volume 116, and a bleed air duct 120.

In operation, the compressor portion 102 compresses air in an airflowpath indicated by the arrow 101. The airflow path flows into thediffuser portion 106. The diffuser portion reduces the velocity of thecompressed air by increasing the cross-sectional area of the airflowpath. A portion of the compressed air contacts the impingement sleeve110 and flows along the outer surface of the transition piece 108. Theflow of air along the outer surface of the transition piece 108 coolsthe transition piece 108, and enters the head end volume 116.

The bleed air duct 120 draws bleed air from the airflow path via avortex cavity 122. The bled air increases the effectiveness of thediffuser portion 106 by improving the diffuser pressure-recoverycoefficient. The arrangement of the impingement sleeve 110 in theairflow path induces a low driving pressure that improves the ducting ofbleed air into the bleed air duct 120. The bleed air is routed into thehead end volume 116.

The bleed air and the transition piece 108 cooling air mix in the headend volume 116 and enters the fuel nozzles 118. The air mixes with fueland is discharged from the fuel nozzles 118 into the combustion chamber124 where the fuel air mixture is ignited. The arrangement provides thebleed extraction for boundary layer control and efficiency, low pressureloss operation of the diffuser, and routes the bleed air to the fuelnozzles upstream of the first turbine rotor stage. This arrangement andsequence allows the air to be used for premixing with fuel, lowering theemission of nitrogen oxides, and also avoids injection of the bleed airdownstream of the first turbine rotor, increasing output and efficiencyrelative to a downstream air return arrangement.

Routing the bleed air to the fuel nozzles increases the efficiency ofthe engine and decreases undesirable emissions since the air removedfrom the air stream is used in the combustion of fuel. Previous systemsand methods routed the bleed air down stream from the compressor, whichmay reduce output performance and efficiency and increase exhaustpollution levels.

The illustrated embodiment shows the bleed air duct 120 routed throughthe outer casing 104, the mount 112, and the bracket 114. The use of theouter casing 104, the mount 112, and the bracket 114 to define the bleedair duct 120 decreases the packaging area in the gas turbine 100.Alternate embodiments may include a second vortex cavity 126 that isoperative to draw additional bleed air, and rout the bleed air to thebleed air duct 120. Other alternate embodiments may include any numberof vortex cavities that draw bleed air into the bleed air duct 120.

FIG. 2 illustrates a front partially-cut away view of the gas turbineengine along the line A-A (of FIG. 1). FIG. 2 shows the bleed air duct120 partially defined by the bracket 114 having a Y-shape. The bracket114 supports the transition piece 108. Other embodiments may include abracket having a single bleed air duct 120 path as opposed to a Y-shapedduct.

While the invention has been described in detail in connection with onlya limited number of embodiments, it should be readily understood thatthe invention is not limited to such disclosed embodiments. Rather, theinvention can be modified to incorporate any number of variations,alterations, substitutions or equivalent arrangements not heretoforedescribed, but which are commensurate with the spirit and scope of theinvention. Additionally, while various embodiments of the invention havebeen described, it is to be understood that aspects of the invention mayinclude only some of the described embodiments. Accordingly, theinvention is not to be seen as limited by the foregoing description, butis only limited by the scope of the appended claims.

What is claimed is:
 1. A gas turbine system comprising: a compressoroperative to output an airstream; a diffuser having an inlet to receivethe airstream and an outlet to output the airstream, the outlet havingan area larger than the inlet to diffuse the airstream; a fuel nozzleoperative to receive fuel and emit the fuel in a combustor; at least onebleed duct having a first and a second inlet between the compressor andthe outlet of the diffuser, the second inlet located upstream of thefirst inlet, and the at least one bleed duct operative to direct bleedair from downstream of the compressor directly to the fuel nozzle; afirst vortex cavity at the first inlet of the bleed duct operative todraw compressed air from the at least one airstream in the diffuser; anda second vortex cavity at the second inlet, the first and second vortexcavities arranged such that bleed air from each of the first and secondvortex cavities flows into the same bleed duct prior to flowing to thefuel nozzle.
 2. The system of claim 1, wherein the system furthercomprises an outer casing that defines the diffuser, and the firstvortex cavity is a recess in the outer casing.
 3. The system of claim 1,wherein the first vortex cavity is located between the inlet and theoutlet of the diffuser.
 4. The system of claim 1, wherein the systemfurther comprises an outer casing that defines the diffuser, and the atleast one bleed duct passes through the outer casing.
 5. The system ofclaim 1, wherein the fuel nozzle is partially disposed in a cavity, thecavity operative to receive the bleed air and rout bleed air to the fuelnozzle.
 6. The system of claim 1, wherein the system further comprisesan impingement sleeve operative to induce a driving pressure on theairstream.
 7. The system of claim 1, wherein the first vortex cavity islocated up stream of the diffuser.
 8. The system of claim 1, wherein thesystem further comprises an outer casing that defines the diffuser, theat least one bleed duct passes through the outer casing to the fuelnozzle, and the airstream flows out from the outlet of the diffuser tothe fuel nozzle.
 9. The gas turbine system of claim 1, wherein the bleedduct has a volume less than a volume of the diffuser.
 10. The gasturbine system of claim 1, wherein the diffuser has a substantiallyannular cross-sectional shape, and the bleed duct comprises a channelextending substantially radially from the substantially annularcross-sectional shape of the diffuser.
 11. The gas turbine system ofclaim 1, wherein the airstream output from the diffuser is directed tothe fuel nozzle to be mixed with fuel and combusted in the combustor.12. The gas turbine system of claim 1, wherein the at least one bleedduct and the diffuser are configured such that bleed air output from theat least one bleed duct is mixed with the airstream output from thediffuser upstream of the fuel nozzle.
 13. A method for routing bleed aircomprising: outputting an airstream of compressed air from a compressor;diffusing the airstream in a diffuser; directing the bleed air through ableed duct directly to a fuel nozzle, the bleed duct comprising a firstinlet and a second inlet located upstream of the first inlet; drawing,with a first vortex cavity located at the first inlet upstream from anoutlet of the diffuser, bleed air from the airstream downstream from thecompressor and upstream from the outlet of the diffuser; drawing, with asecond vortex cavity located at the second inlet of the bleed duct,bleed air from the airstream downstream from the compressor; andcombining the bleed air from the first vortex cavity with the bleed airfrom the second vortex cavity.
 14. The method of claim 13, wherein thebleed air is drawn from the airstream by the first vortex cavity locatedin the diffuser.
 15. The method of claim 13, wherein a driving pressureis induced on the air stream by an impingement sleeve.
 16. The method ofclaim 13, further comprising: directing the airstream from the diffuseralong an impingement sleeve to the fuel nozzle.